The disclosed embodiments generally pertain to one or more methods of forming a ceramic matrix composite aircraft engine component. More particularly, but not by way of limitation, present embodiments relate to a method of forming a ceramic matrix composite turbine blade squealer tip and pressure side flare.
A typical gas turbine engine generally possesses a forward end and an aft end with its several core or propulsion components positioned axially therebetween. An air inlet or intake is at a forward end of the gas turbine engine. Moving toward the aft end, in order, the intake is followed by a compressor, a combustion chamber and a turbine toward the aft end of the gas turbine engine. It will be readily apparent from those skilled in the art that additional components may also be included in the gas turbine engine, such as, for example, low-pressure and high-pressure compressors, and high-pressure and low-pressure turbines. This, however, is not an exhaustive list. A gas turbine engine also typically has an internal shaft axially disposed along a center longitudinal axis of the gas turbine engine. The internal shaft is connected to both the turbine and the air compressor, such that the turbine provides a rotational input to the air compressor to drive the compressor blades.
In operation, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases which flow downstream through turbine stages. These turbine stages extract energy from the combustion gases. A high pressure turbine first receives the hot combustion gases from the combustor and includes a stator nozzle assembly directing the combustion gases downstream through a row of high pressure turbine rotor blades extending radially outwardly from a supporting rotor disk. In a two stage turbine, a second stage stator nozzle assembly is positioned downstream of the first stage blades followed in turn by a row of second stage rotor blades extending radially outwardly from a second supporting rotor disk. The turbine converts the combustion gas energy to mechanical energy wherein each set of stator vanes turns and accelerates the combustion gases to engage an adjacent row of rotating turbine blades.
In the formation of components for aircraft and aircraft engines, such as, for non-limiting example, turbine blades, a squealer tip may be formed, which is an open cavity at the tip end of the turbine blade. Additionally, the pressure side and/or the suction side tip of the blade may be flared as both of these features improve turbine blade efficiency. In forming turbine blades, the airfoil total thickness may be either constant or increasing from tip to root or hub so that there are no part-span plies, which would be unsupported from the root. However, in forming a tip flare, additional material is required to provide the tip flare which violates the constraint pertaining to airfoil total thickness.
Additionally, when forming a squealer tip in a CMC blade, one exemplary method of forming such squealer tip is that the cavity would need to be machined after the part is formed. However, these additional machining processes that are required to form the cavity require additional costs to form the part. This machining after the part is formed is inefficient for manufacturing and is a cost adder to the production of the turbine blades.
As may be seen by the foregoing, improving the manufacture of gas turbine engine components may be beneficial. For example, it may be beneficial to reduce tip flow leakage, or increase turbine efficiency, or improve tip cooling or any combination of these. Moreover, it may be beneficial to form a ceramic matrix composite turbine blade which includes a squealer tip and tip flare while meeting the desired constraints for turbine blade architecture.
The information included in this Background section of the specification, including any references cited herein and any description or discussion thereof, is included for technical reference purposes only and is not to be regarded subject matter by which the scope of the instant embodiments are to be bound.